JP4912522B2 - Ceramic turbine nozzle - Google Patents

Ceramic turbine nozzle Download PDF

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Publication number
JP4912522B2
JP4912522B2 JP2000220197A JP2000220197A JP4912522B2 JP 4912522 B2 JP4912522 B2 JP 4912522B2 JP 2000220197 A JP2000220197 A JP 2000220197A JP 2000220197 A JP2000220197 A JP 2000220197A JP 4912522 B2 JP4912522 B2 JP 4912522B2
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ceramic
band
rear portion
bands
blade
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JP2001090505A5 (en
JP2001090505A (en
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アンジェロ・ファン・コスチャー
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General Electric Co
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General Electric Co
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/146—Shape, i.e. outer, aerodynamic form of blades with tandem configuration, split blades or slotted blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/284—Selection of ceramic materials
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
    • F05D2300/603—Composites; e.g. fibre-reinforced

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Ceramic Engineering (AREA)
  • Materials Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Description

【０００１】
【発明の属する技術分野】
本発明は概してガスタービンエンジンに関し、より具体的にはそのタービンノズルに関する。
【０００２】
【従来の技術】
ガスタービンエンジンでは、空気がコンプレッサーで加圧され、燃焼器で燃料と混合されそして点火されて高温の燃焼ガスを発生し、この高温の燃焼ガスが下流のタービン中に流れ込み、タービンでガスからエネルギーを抽出する。タービンは、一体の外側及び内側バンドにより支持された複数の円周方向に離れて配置されたノズル羽根を有するタービンノズルを含む。高圧タービンノズルは最初に、最も高温の燃焼ガスを燃焼器から受け、支持ディスクから半径方向外方に延びる複数の円周方向に離れて配置された動翼を有するタービンロータにそれらのガスを流す。
【０００３】
全体的なエンジン効率は燃焼ガスの温度に直接関係し、燃焼ガスの温度はガスによって熱せられる様々なタービン構成部品を保護するために制限されなければならない。高圧タービンノズルは、適当な耐用寿命を考えて燃焼器からの高温燃焼ガスに耐えなければならない。これは一般に高温で強度を保持する超合金材料を用いることと、コンプレッサの空気の一部をタービンノズルにおける冷却媒体として使用する目的で分流することとによって実現される。
【０００４】
超合金の強度には限度があり、分流されるコンプレッサの空気はエンジンの全体的な効率を減ずる。従って、エンジンの効率は、適当な超合金の利用可能性及びタービンノズルを冷却するためにコンプレッサの空気を分流する必要性により事実上制限される。
【０００５】
タービンノズルの温度性能をさらに増し、タービンノズルのために分流される冷却空気の使用を削減するようにタービンノズルを改善するために、セラミック系材料が考えられている。しかしながら、この目的のために利用可能な従来のセラミック材料は、延性がほとんどなくその耐用年数を制限するその破損損傷を防ぐためには特別な取り付け構造を必要とする。
【０００６】
ノズルは、三次元の空力荷重及びそれを通しての温度勾配の影響下にある羽根の環状の組立体であるため、タービンノズルの設計はさらに複雑になっている。タービンノズルは、運転中に膨張したり収縮したりし、その結果熱応力を生じる。
【０００７】
モノリシック・セラミックは簡単に成形できるが、その一体的な接合点で比較的に脆弱である。セラミック母材の複合材（CMC）は、機械的強度を増大させる意図でセラミック母材にセラミック繊維を採り入れる。繊維は結合母材に強度を与える。しかしながら、セラミック繊維はほとんど延性を有しないため、曲げたりまたタービンノズルのような複雑な三次元構成部品において要求される移行に対応するには不十分な能力しか備わっていない。
【０００８】
従って、ガスタービンエンジンの厳しい環境に耐えるためにセラミックから形成される改良されたタービンノズルを供給することが望まれる。
【０００９】
【発明の開示】
タービンノズルはセラミックの外側及び内側バンドを含み、セラミック羽根前部分がそれに一体的に結合されている。セラミック羽根後部分は両端でバンドの相補的ソケットに嵌合して（とじ込められて）いる。
【００１０】
【発明の実施の形態】
好ましいかつ例示的な実施形態に従って、本発明を、そのさらなる目的と利点とともに、添付の図面に関連してなされる以下の詳細な記述により具体的に説明する。
【００１１】
図１に示されるのは、ガスタービンエンジン用の環状の高圧タービンノズル１０の１部であり、高温の燃焼ガス１２をノズルに排出するガスタービンエンジンの燃焼器の下流に位置している。ノズルはセラミックのアーチ形になった外側及び内側バンド１４，１６を含む。バンドはリングのセグメントであってもよいしあるいは要求があれば連続したリングであってもよい。
【００１２】
円周方向に離れて配置された複数のセラミック羽根１８が外側及び内側バンドの間に取り付けられるが、図１には、例示的なノズルセグメントを示すために２枚の羽根が図示されている。それぞれの羽根は図２により詳細に示されるような適当な翼形状を有し、軸方向に対向する前縁１８ａ及び後縁１８ｂを含み、それらが円周方向即ち横方向に対向する正圧側面１８ｃ及び負圧側面１８ｄを互いに結合する。従来の慣行に従って燃焼ガスの向きを変える必要から、正圧側面１８ｃは普通凹状であり負圧側面１８ｄは普通凸状である。
【００１３】
実際に役立つセラミック・タービンノズルを構成するためには、それぞれの羽根１８が１組の相補的な羽根部分によって画成される。羽根前部２０は、構造的強度を備えるために単体のすなわち一体型の組立体として半径方向の両端でバンド１４，１６のうちの対応するバンドと一体的に結合される。羽根後部２２は、バンド１４，１６のうちそれぞれのバンドにある相補的ソケット２４に嵌合される対向する半径方向外側及び内側端２２ａを有する。
【００１４】
この構成では、羽根部２０，２２は両方とも、低い延性のセラミックが用いられているにもかかわらず、運転中に適当な強度を実現するようにタービンノズルに必要な複雑な三次元構成においてセラミックで形成され得る。
【００１５】
図１及び図２に示される好ましい実施形態においては、各羽根前部分２０は、環状のタービンノズルが調整された方向性強度を有し、そして一体型のバンド１４，１６と強力に結合するために従来のセラミック母材の複合材（CMC）を用いて形成され得る。これらの図に概略的に示されるように、前部分２０は適当なセラミック母材２０ｂ中にセラミック繊維編組２０ａを含むのが望ましい。従来のセラミック母材の複合材が利用可能であるが、炭化ケイ素母材（ＳｉＣ）中に炭化ケイ素繊維（SｉＣ）を含むものとすることができる。その繊維と母材は、最初は、一般的に柔軟性がある素地の状態の適当な母材に含まれており、処理されつまり硬化して最終的なセラミック状態になる。
【００１６】
図３に示される好ましい実施形態においては、セラミック繊維編組２０ａは最初は途切れることのない管状の連続した繊維の形状をしている。その管は羽根前部分の所望の輪郭を有する適当な工具類を用いて簡単に成形される。外側及び内側バンド１４，１６は、強度を増大するため前部分編組２０ａと共に適当に積層され得るCMC積層体１４ａ，１６ａの形状であることが望ましい。
【００１７】
もっと具体的に言えば、図３に示される編組管２０ａは、バンド積層体と共に積層するための一体的移行部を備えた張り広げられたつまりキノコ状をした両端２０ｃの形状にスリットを入れられた長手方向両端を有することが望ましい。前部２０及びバンド１４，１６は両方とも、望ましくは同一のセラミック母材に同一のセラミック繊維を用いたCMCで形成されることが望ましい。
【００１８】
編組管２０ａは、バンド間に必要な半径方向の広がりをもって完成した翼形部の前縁部を形成するように構成され、そして、張り広げられた端２０ｃは部分的にそれらのバンドを形成するように対応するバンドに沿って向け直され得る。円周方向に隣接する前部分の張り広げられた端は、バンドの周囲に沿ってお互いに隣接し、バンドはその他は必要なバンドの形状になるようにCMCテープまたは織物積層体を用いて完成される。処理されすなわち硬化すると、素地の前部分及びバンドはその最終セラミック状態で堅まり、これらの構成部品の一体構造組立体となる。
【００１９】
この組立体の特別な利点は、羽根前部分２０がそれの織り合わされた繊維により最高の強度性能を有する編組管で形成されることである。それらの繊維はセラミックであるので、ほとんど延性を有しないがそれでも張り広げられた端２０ｃのあるなしにかかわらずバンドと一体的に形成され得る。
【００２０】
図３に示されるように、編組２０ａ中のセラミック繊維は、前部分とバンドの間に形成される最終のコーナー丸み部分にわたって羽根前部分から対向する外側及び内側バンドへ傾斜角度Aで移行するのが望ましい。比較的に剛性のあるセラミック繊維による羽根とバンドの交差部での結果として生じる丸みを最小にするためには、好ましい実施形態では傾斜角度は約４５°までならよい。
【００２１】
従って、張り広げられた編組端２０ｃは、そこに積層される外側及び内側バンド１４，１６との一体構造を提供し、タービンノズルに対して主たる強度を提供する。編組端はバンド積層体とクロスステッチにするかまたはバンド積層体と重ね合せにすることができる。羽根前部分及びバンド中のセラミック繊維は、運転中に受ける三次元の荷重及び温度差に対して要求される方向でのノズル強度を最大にする方向に優先的に向けることができる。
【００２２】
図２で初めに示されるように、それぞれの羽根１８は、比較的に大きい半径の前縁１８ａ及び比較的に薄い半径の後縁１８ｂを備える空気力学的な三日月形の輪郭を有する。後縁の半径は、ノズルの空力性能を最大にするために必要な一般に約１０ミルである。そのような薄い後縁は、セラミック構造に固有の制約を考慮すると複合材のタービンノズルの設計をさらに複雑なものにする。セラミック繊維はほとんど延性を備えていないので、それらの繊維を薄い後縁に必要な小さな半径の周りに曲げるのは一般的に不可能である。さらに、CMC複合材の層の厚さもまた一般的には薄い羽根後縁の厚さよりは大きい。
【００２３】
羽根は燃焼ガスを流すように構成されるていので、羽根は運転中にガス圧によって高い負荷を受け、またガスの高温に曝され温度差による熱膨張及び収縮を引き起こす。そして、羽根後縁は比較的に薄いので、それの冷却を行える空間を設ける余地はほとんどない。
【００２４】
従って、図１から図３までに示される好ましい実施形態においては、各羽根後部分２２はその中に強化セラミック繊維のないモノリシック・セラミックから成る。モノリシック・セラミックは窒化ケイ素（Sｉ3Ｎ4）のような従来型のものである。羽根後部分２２は高靭性モノリシック・セラミックで形成されるのが望ましいけれども、それらは一般に前部分２０に見られる向きとは異なる向きにその中の強化セラミック繊維を備えたセラミック複合材で形成してもよい。
【００２５】
例えば、前部分２０中の繊維は傾斜方向の角度Ａに向いているのが望ましいのに対して、後部分22に用いられる繊維は、後縁の半径方向の強度を増すために後部分の両端の間で半径方向に延びるていのが望ましいであろう。後部分における繊維の好適な半径方向の向きを考慮すると、或いは後部分の他の態様のモノリシック構造を考慮すると、後部分の外側及び内側バンドへの特別な取付けがノズル組立体及びその強度を補っている。
【００２６】
上記のように、羽根後部分２２は一体になった前部分及びバンドから分離し異なるものであることが望ましい。前部分及びバンドにより画成される構造枠を用いて、個々の後部分をそれらが対応する前部分に隣接する位置に機械的に嵌合し、個々の空力羽根を完成するのが有利である。
【００２７】
図１及び図３に示されるように、各後部分の半径方向の外側及び内側両端２２ａには、後部分から延び出る軸方向に細長い支持キーを形成するのが望ましい。その支持キー２２ａは、対応する外側及び内側バンドに形成される相補的な座すなわちソケット２４に簡単に嵌合され、それぞれの後部分を外側及び内側バンドの間に保持し羽根にかかるトルクをバンドに搬送する。この構造では、後部分は、それらが嵌合される外側及び内側バンドに対して半径方向に膨張したり収縮したりすることが可能である。そして、後部分にかかる空力トルク負荷は、支持キー２２ａを介して対応するバンドで担持される。
【００２８】
このようにして、CMC羽根前部分２０は、セラミック繊維で強化された外側及び内側バンドと共に構造枠を構成する。また、薄い羽根後部分は空力性能を最大にするために特別の輪郭形状とすることができ、嵌合によってバンドの間に保持することができる。したがって、他の実施形態では、後部分は実行可能なら繊維で強化しているが、モノリシック・セラミックは後部分に選択的に有利に用いることができる。
【００２９】
例えば図２に示される２部分からなる構造においては、望ましくは羽根後部分２２は羽根前部分２０から間隔を置いて配置され、それらの間に小さな間隙２６を設けている。羽根部分２０，２２のどちらか一方又は両方はコンプレッサの抽気空気等の冷却媒体２８をその中に流すために、半径方向に中空にすることができる。各部分はまた間隙内に隠された列になった吐出孔３０を含むこともでき、運転中に間隙の中へ冷却媒体を吐出する。このようにして、冷却媒体がなんらかの適当な方法でその内部冷却するために各羽根部分を貫通して流され、その後冷却媒体は間隙２６の中へ吐出され後部分の外側表面を覆うように下流に流れるにつれて、冷却空気の膜を形成する。
【００３０】
差圧が運転中に各羽根の両側面１８ｃと１８ｄとの間に生じるので、各羽根は図２に示すように間隙２６内で羽根前部分２０と後部分２２の間に配置されるシール３２を含み、そこを流れる流体をシールするのが望ましい。シール３２は、間隙２６を画成する面の相補的な凹陥に嵌装されたセラミックロープのシールのようなどのような適当な構成をしていてもよい。シールは高温の燃焼ガスが間隙２６を介して流れるのを阻止する一方、シールの両横方向側面上を間隙２６を介して冷却媒体２８が吐出するのを可能にしている。
【００３１】
図３は、図１及び図２に示されるセラミック・タービンノズル１０の好ましい製造法の概略を示す。各羽根後部分２２は、例えばモノリシック材料を後部分の所望の構成に成形する等どのような適当な方法ででも予備成形されることが望ましい。
【００３２】
個々のセラミック繊維管２０ａは、その素地の状態で羽根前部分の所望の形状に形成され、対応する後部分２２を補完し、両者によって個々の羽根１８を構成する。各前部分の張り広げられた端２０ｃはその後素地の状態で外側及び内側バンドのセラミック織物と積層される。
【００３３】
このようにして、前部分及びバンドのセラミック構成要素は適当な工具または型枠を使って要求される形に形成されるかモールドされ、個々の予備成形された後部分２２がそれに組み合わされる。したがって、後部分は組立工程においてバンドの間であって対応する前部分の後方に嵌合される。
【００３４】
次ぎに、素地のバンドと前部分は従来の方法で処理されるかまたは硬化され、後部分がその中に機械的に嵌合された硬化したセラミックノズルを形成する。
【００３５】
この好ましい構成では、羽根後部分２２は、強化セラミック繊維をもたないモノリシック・セラミックのような前もって硬化処理されたセラミックであることが望ましい。そして、羽根前部分２０及びバンド１４，１６は、その中に強化セラミック繊維を有するセラミック母材の複合材構造であり、構造的に一体のものとなりまた組立体全体に強度を与える。
【００３６】
この構造では、管編組２０ａの強度上の利点が羽根前部分をバンドと一体化するために用いられ、羽根後部分２２はバンドに機械的に保持されるか又は嵌合される。後部分は軸方向及び円周方向にバンドに保持されるが、支持ソケット２４内でバンド間で半径方向に自由に膨張したり収縮したりする。
【００３７】
セラミック母材の複合材及びモノリシック・セラミックのそれぞれの利点を選択的に用いて、タービンノズルをその一体性と耐久性を最大にするように構成する。羽根前部分及び後部分２０、２２の相対的な大きさは、CMC及びモノリシック・セラミック材料の製造能力の要求に合わせて調整することができる。
【００３８】
本明細書では本発明の好ましい例示的な実施形態と思料するものについて説明してきたが、本発明のその他の形態は本明細書の教示内容から当業者には自明であり、本発明の技術的思想及び技術的範囲に属するかかる形態すべてが特許請求の範囲で保護されることを望むものである。
【００３９】
従って、特許による保護を望むのは、請求項に規定され特徴付けられた発明である。
【図面の簡単な説明】
【図１】 本発明の例示的な実施形態による環状のセラミック・タービンノズルのセグメントの等角図。
【図２】 図１に示されたセラミック羽根の１つを線２−２に沿って見た半径方向断面図。
【図３】 図１及び図２に示されたセラミック・タービンノズルの例示的な製造法のフローチャート図。[0001]
BACKGROUND OF THE INVENTION
The present invention relates generally to gas turbine engines, and more specifically to the turbine nozzle.
[0002]
[Prior art]
In a gas turbine engine, air is pressurized with a compressor, mixed with fuel in a combustor and ignited to generate hot combustion gases that flow into the downstream turbine, where the energy from the gases is absorbed by the turbine. To extract. The turbine includes a turbine nozzle having a plurality of circumferentially spaced nozzle vanes supported by integral outer and inner bands. The high pressure turbine nozzle initially receives the hottest combustion gases from the combustor and flows them through a turbine rotor having a plurality of circumferentially spaced blades extending radially outward from the support disk. .
[0003]
Overall engine efficiency is directly related to the temperature of the combustion gas, which must be limited to protect the various turbine components that are heated by the gas. The high pressure turbine nozzle must withstand the high temperature combustion gases from the combustor for an appropriate service life. This is generally achieved by using a superalloy material that retains strength at high temperatures and diverting a portion of the compressor air for the purpose of use as a cooling medium in the turbine nozzle.
[0004]
Superalloy strength is limited, and the shunting compressor air reduces the overall efficiency of the engine. Thus, engine efficiency is effectively limited by the availability of suitable superalloys and the need to divert compressor air to cool the turbine nozzle.
[0005]
Ceramic-based materials are being considered to improve the turbine nozzle to further increase the temperature performance of the turbine nozzle and reduce the use of cooling air diverted for the turbine nozzle. However, the conventional ceramic materials available for this purpose require special mounting structures to prevent their breakage damage, which has little ductility and limits its useful life.
[0006]
Turbine nozzle design is further complicated because the nozzle is an annular assembly of blades that is subject to three-dimensional aerodynamic loads and temperature gradients therethrough. Turbine nozzles expand and contract during operation, resulting in thermal stress.
[0007]
Monolithic ceramics can be easily formed, but are relatively fragile at their integral joints. Ceramic matrix composites (CMC) incorporate ceramic fibers into the ceramic matrix with the intention of increasing mechanical strength. The fiber provides strength to the bond matrix. However, because ceramic fibers have little ductility, they have insufficient capacity to bend and accommodate the transitions required in complex three-dimensional components such as turbine nozzles.
[0008]
Accordingly, it would be desirable to provide an improved turbine nozzle formed from ceramic to withstand the harsh environment of a gas turbine engine.
[0009]
DISCLOSURE OF THE INVENTION
The turbine nozzle includes ceramic outer and inner bands with a ceramic blade front portion integrally coupled thereto. The rear part of the ceramic blade is fitted (captured) into the complementary socket of the band at both ends.
[0010]
DETAILED DESCRIPTION OF THE INVENTION
In accordance with the preferred and exemplary embodiments, the invention, together with further objects and advantages thereof, will be specifically described by the following detailed description taken in conjunction with the accompanying drawings.
[0011]
Illustrated in FIG. 1 is a portion of an annular high pressure turbine nozzle 10 for a gas turbine engine, located downstream of a combustor of a gas turbine engine that discharges hot combustion gas 12 to the nozzle. The nozzle includes ceramic arched outer and inner bands 14,16. The band may be a segment of the ring or a continuous ring if required.
[0012]
Although a plurality of circumferentially spaced ceramic blades 18 are mounted between the outer and inner bands, two blades are shown in FIG. 1 to illustrate an exemplary nozzle segment. Each vane has a suitable airfoil shape as shown in more detail in FIG. 2 and includes an axially opposed leading edge 18a and a trailing edge 18b, which are circumferentially or laterally opposed pressure sides. 18c and suction side 18d are coupled together. The pressure side 18c is normally concave and the suction side 18d is normally convex because it is necessary to change the direction of the combustion gas in accordance with conventional practice.
[0013]
In order to construct a practical ceramic turbine nozzle, each blade 18 is defined by a set of complementary blade portions. The vane front portion 20 is integrally joined with the corresponding one of the bands 14, 16 at both radial ends as a unitary or unitary assembly to provide structural strength. The vane rear portion 22 has opposing radially outer and inner ends 22a that fit into complementary sockets 24 in each of the bands 14,16.
[0014]
In this configuration, both vanes 20 and 22 are ceramic in the complex three-dimensional configuration required for the turbine nozzle to provide adequate strength during operation, despite the use of low ductility ceramic. Can be formed.
[0015]
In the preferred embodiment shown in FIGS. 1 and 2, each vane front portion 20 has an annular turbine nozzle with adjusted directional strength and a strong coupling with the integral bands 14,16. Or a conventional ceramic matrix composite (CMC). As schematically shown in these figures, the front portion 20 preferably includes a ceramic fiber braid 20a in a suitable ceramic matrix 20b. Conventional ceramic matrix composites can be used, but silicon carbide fibers (SiC) can be included in the silicon carbide matrix (SiC). The fibers and matrix are initially contained in a suitable matrix in a generally flexible substrate state and are processed or cured to a final ceramic state.
[0016]
In the preferred embodiment shown in FIG. 3, the ceramic fiber braid 20a is initially in the form of an unbroken tubular continuous fiber. The tube is simply formed using suitable tools having the desired contour of the vane front portion. The outer and inner bands 14, 16 are preferably in the form of CMC laminates 14a, 16a that can be suitably laminated with the front braid 20a to increase strength.
[0017]
More specifically, the braided tube 20a shown in FIG. 3 is slit into the shape of the stretched or mushroomed ends 20c with integral transitions for lamination with the band laminate. It is desirable to have both longitudinal ends. Both the front portion 20 and the bands 14 and 16 are preferably formed of CMC using the same ceramic fiber in the same ceramic base material.
[0018]
The braided tube 20a is configured to form the leading edge of the completed airfoil with the required radial extent between the bands, and the stretched end 20c partially forms those bands. Can be redirected along the corresponding band. The stretched edges of the circumferentially adjacent front part are adjacent to each other along the circumference of the band, and the band is completed with CMC tape or woven laminate so that the others are in the required band shape Is done. When processed or hardened, the front part of the substrate and the band are hardened in their final ceramic state, resulting in a unitary assembly of these components.
[0019]
A particular advantage of this assembly is that the vane front portion 20 is formed of a braided tube having the highest strength performance due to its interwoven fibers. Since these fibers are ceramic, they have little ductility but can still be integrally formed with the band with or without stretched ends 20c.
[0020]
As shown in FIG. 3, the ceramic fibers in the braid 20a transition at an inclination angle A from the blade front portion to the opposite outer and inner bands over the final corner rounded portion formed between the front portion and the band. Is desirable. In order to minimize the resulting roundness at the blade-band intersection due to the relatively stiff ceramic fibers, in preferred embodiments the tilt angle may be up to about 45 °.
[0021]
Thus, the stretched braid end 20c provides a unitary structure with the outer and inner bands 14, 16 laminated thereon and provides the main strength to the turbine nozzle. The braided end can be cross-stitched with the band laminate or can be overlapped with the band laminate. The ceramic fibers in the vane front part and the band can be preferentially directed in a direction that maximizes the nozzle strength in the direction required for the three-dimensional loads and temperature differences experienced during operation.
[0022]
As initially shown in FIG. 2, each vane 18 has an aerodynamic crescent profile with a relatively large radius leading edge 18a and a relatively thin radius trailing edge 18b. The trailing edge radius is typically about 10 mils required to maximize the aerodynamic performance of the nozzle. Such a thin trailing edge further complicates the design of a composite turbine nozzle in view of the limitations inherent in ceramic structures. Because ceramic fibers have little ductility, it is generally impossible to bend them around the small radius required for a thin trailing edge. Furthermore, the thickness of the CMC composite layer is also generally greater than the thickness of the thin blade trailing edge.
[0023]
Since the vanes are configured to flow combustion gas, the vanes are subjected to high loads due to gas pressure during operation, and are exposed to high gas temperatures, causing thermal expansion and contraction due to temperature differences. Since the trailing edge of the blade is relatively thin, there is little room for providing a space for cooling the blade.
[0024]
Accordingly, in the preferred embodiment shown in FIGS. 1-3, each vane rear portion 22 is comprised of a monolithic ceramic without reinforced ceramic fibers therein. Monolithic ceramics are conventional, such as silicon nitride (Si 3 N 4 ). Although the back vane portion 22 is preferably formed of a high toughness monolithic ceramic, they are generally formed of a ceramic composite with reinforcing ceramic fibers therein in a different orientation than that found in the front portion 20. Also good.
[0025]
For example, it is desirable for the fibers in the front portion 20 to be oriented at an angle A in the tilt direction, whereas the fibers used in the rear portion 22 are both ends of the rear portion to increase the radial strength of the trailing edge. It would be desirable to extend radially between the two. Considering the preferred radial orientation of the fibers in the rear part, or considering other aspects of the monolithic structure of the rear part, special attachments to the outer and inner bands of the rear part supplement the nozzle assembly and its strength. ing.
[0026]
As mentioned above, the vane rear portion 22 is preferably separate and distinct from the integrated front portion and band. With the structural frame defined by the front part and the band, it is advantageous to mechanically fit the individual rear parts in positions adjacent to the corresponding front part to complete the individual aerodynamic blades. .
[0027]
As shown in FIGS. 1 and 3, it is desirable to form axially elongated support keys extending from the rear portion at the radially outer and inner ends 22a of each rear portion. The support key 22a is simply fitted into complementary seats or sockets 24 formed on the corresponding outer and inner bands, holding the respective rear portions between the outer and inner bands and providing torque to the vanes. Transport to. With this construction, the rear portions can expand and contract radially relative to the outer and inner bands into which they are fitted. The aerodynamic torque load applied to the rear portion is carried by the corresponding band via the support key 22a.
[0028]
In this way, the CMC vane front portion 20 constitutes a structural frame with outer and inner bands reinforced with ceramic fibers. In addition, the thin rear portion of the blade can be specially contoured to maximize aerodynamic performance and can be held between the bands by fitting. Thus, in other embodiments, the back portion is reinforced with fibers if practicable, but monolithic ceramics can be selectively used advantageously in the back portion.
[0029]
For example, in the two-part construction shown in FIG. 2, the trailing blade portion 22 is preferably spaced from the leading blade portion 20 with a small gap 26 therebetween. Either or both of the vane portions 20, 22 can be radially hollow to allow a cooling medium 28, such as compressor bleed air, to flow therethrough. Each portion may also include a row of discharge holes 30 hidden in the gap to discharge the cooling medium into the gap during operation. In this way, the cooling medium is flowed through each vane portion for internal cooling in any suitable manner, after which the cooling medium is discharged into the gap 26 and downstream so as to cover the outer surface of the rear portion. As a result, a film of cooling air is formed.
[0030]
Since differential pressure occurs between the side surfaces 18c and 18d of each blade during operation, each blade is a seal 32 disposed between the blade front portion 20 and the rear portion 22 within the gap 26 as shown in FIG. It is desirable to seal the fluid flowing therethrough. The seal 32 may be of any suitable configuration, such as a ceramic rope seal fitted into a complementary recess in the surface defining the gap 26. The seal prevents hot combustion gases from flowing through the gap 26 while allowing the cooling medium 28 to discharge through the gap 26 on both lateral sides of the seal.
[0031]
FIG. 3 outlines a preferred method of manufacturing the ceramic turbine nozzle 10 shown in FIGS. Each vane rear portion 22 is preferably preformed in any suitable manner, such as by molding a monolithic material into the desired configuration of the rear portion.
[0032]
The individual ceramic fiber tubes 20a are formed in a desired shape of the front part of the blade in the state of the base, complement the corresponding rear part 22, and the individual blade 18 is constituted by both. The stretched end 20c of each front portion is then laminated with ceramic fabric of the outer and inner bands in a green state.
[0033]
In this way, the front and band ceramic components are formed or molded into the required shape using a suitable tool or formwork and the individual preformed rear portions 22 are combined therewith. Thus, the rear part is fitted between the bands and behind the corresponding front part in the assembly process.
[0034]
The green band and front portion are then processed or cured in a conventional manner to form a hardened ceramic nozzle with the rear portion mechanically fitted therein.
[0035]
In this preferred configuration, the vane rear portion 22 is preferably a pre-cured ceramic, such as a monolithic ceramic without reinforcing ceramic fibers. The blade front portion 20 and the bands 14 and 16 have a composite structure of a ceramic base material having reinforcing ceramic fibers therein, and are structurally integrated and give strength to the entire assembly.
[0036]
In this construction, the strength advantage of the tube braid 20a is used to integrate the vane front portion with the band, and the vane rear portion 22 is mechanically held or fitted to the band. The rear portion is held by the band in the axial direction and the circumferential direction, but freely expands and contracts between the bands in the support socket 24 in the radial direction.
[0037]
The advantages of each of the ceramic matrix composite and the monolithic ceramic are selectively used to configure the turbine nozzle to maximize its integrity and durability. The relative size of the vane front and rear portions 20,22 can be tailored to meet the production capacity requirements of CMC and monolithic ceramic materials.
[0038]
While this specification has described what is considered to be a preferred exemplary embodiment of the present invention, other forms of the invention will be apparent to those skilled in the art from the teachings herein and the technical scope of the invention. All such forms belonging to the spirit and technical scope are desired to be protected by the claims.
[0039]
Accordingly, what is desired to be protected by patent is an invention as defined and characterized in the claims.
[Brief description of the drawings]
FIG. 1 is an isometric view of a segment of an annular ceramic turbine nozzle according to an exemplary embodiment of the present invention.
FIG. 2 is a radial cross-sectional view of one of the ceramic blades shown in FIG. 1 taken along line 2-2.
FIG. 3 is a flow chart diagram of an exemplary method of manufacturing the ceramic turbine nozzle shown in FIGS. 1 and 2;

Claims (6)

セラミックの外側及び内側バンド（14,16）と、
両端で前記バンドと一体に結合されたセラミック羽根前部分（20）と、
前記バンドの相補的ソケット（24）に嵌合される両端（22a）を有するセラミック羽根後部分（22）と
を備え、
前記羽根前部分（20）がセラミック母材の複合材からなり、前記羽根前部分がセラミック母材（20b）中にセラミック繊維編組（20a）を含み、
前記編組（20a）が、前記バンドに積層された張り広げられた両端（20c）を有する管からなる
ことを特徴とする、タービンノズル（10）。Ceramic outer and inner bands (14,16);
Ceramic blade front part (20) integrally joined to the band at both ends;
A ceramic vane rear portion (22) having opposite ends (22a) fitted to the complementary sockets (24) of the band;
The vane front portion (20) is made from a composite material of ceramic matrix, seen including a ceramic fiber braid (20a) said blade leading portion in ceramic matrix (20b),
The turbine nozzle (10), wherein the braid (20a) comprises a tube having stretched ends (20c) laminated to the band . セラミックの外側及び内側バンドと（14,16）、両端で前記バンドに一体に結合されたセラミック母材の複合材の羽根前部分（20）と、前記バンドの相補的ソケット（24）に嵌合される両端（22a）を有するモノリシック・セラミック羽根後部分（22）とを備え、
前記羽根前部分が、セラミック母材（20b）中に前記バンドに積層される張り広げられた両端（20c）を有するセラミック繊維管状編組（20a）をさらに備えてなることを特徴とする、タービンノズル（10）。Mates with ceramic outer and inner bands (14, 16), ceramic base composite vane front part (20) integrally bonded to the band at both ends, and complementary socket (24) of the band A monolithic ceramic blade rear portion (22) having both ends (22a) to be
A turbine nozzle, wherein the blade front portion further comprises a ceramic fiber tubular braid (20a) having stretched ends (20c) laminated to the band in a ceramic base material (20b) (Ten). 前記羽根後部分が、前記相補的ソケットに嵌合されるその両端の支持キーを含み、前記羽根後部分を前記バンド間に保持し、羽根にかかるトルクをバンドに搬送する請求項２に記載のノズル。 3. The wing rear portion includes support keys at both ends thereof fitted to the complementary socket, holds the wing rear portion between the bands, and transfers torque applied to the wing to the band. nozzle. セラミック母材の複合材の素地からなる外側及び内側バンド（14,16）を用意することと、
前記バンドの相補的ソケット（24）に嵌合される両端（22a）を有するセラミック羽根後部分（22）を形成することと、
前記後部分（22）と相補的な関係にある羽根前部分（20）を、セラミック繊維編組（20a）を含むセラミック母材（20b）の複合材の素地から形成することと、
前記素地の前部分を前記素地の外側及び内側バンド（14,16）と積層することと、
前記後部分を前記バンドの間で前記前部分の後側に嵌合することと、
前記素地のバンド及び前記素地の前部分を硬化させ、そこに嵌合される前記後部分とで前記セラミックノズルを形成することと
を含むセラミック・タービンノズル（10）を製造する方法。Providing outer and inner bands (14, 16) made of a composite body of ceramic matrix;
Forming a ceramic vane rear portion (22) having opposite ends (22a) mated to complementary bands (24) of the band;
Forming a front blade portion (20) in a complementary relationship with the rear portion (22) from a composite matrix of a ceramic matrix (20b) comprising a ceramic fiber braid (20a);
Laminating the front portion of the substrate with the outer and inner bands (14, 16) of the substrate;
Fitting the rear portion between the bands to the rear side of the front portion;
A method of manufacturing a ceramic turbine nozzle (10) comprising: curing the base band and a front portion of the base and forming the ceramic nozzle with the rear portion fitted therein. 前記後部分が前もって硬化処理されたセラミックである請求項４に記載の方法。The method of claim 4 , wherein the rear portion is a pre-cured ceramic. 前記後部分がモノリシック・セラミックである請求項５に記載の方法。The method of claim 5 , wherein the rear portion is a monolithic ceramic.
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US6200092B1 (en) 2001-03-13
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